U.S. patent application number 10/059956 was filed with the patent office on 2002-10-10 for method of clinching the top and bottom ends of z-axis fibers into the respective top and bottom surfaces of a composite laminate.
Invention is credited to Garrett, Scott A., Hook, James M., Johnson, David W., Moyers, Steven G..
Application Number | 20020144767 10/059956 |
Document ID | / |
Family ID | 30003940 |
Filed Date | 2002-10-10 |
United States Patent
Application |
20020144767 |
Kind Code |
A1 |
Johnson, David W. ; et
al. |
October 10, 2002 |
Method of clinching the top and bottom ends of Z-axis fibers into
the respective top and bottom surfaces of a composite laminate
Abstract
A method and apparatus for forming an improved pultruded and
clinched Z-axis fiber reinforced composite laminate structure. The
upper and lower skins and the core are pulled automatically through
tooling where the skin material is wetted-out with resin and the
entire composite laminate is preformed in nearly its final
thickness. The preformed composite laminate continues to be pulled
into an automatic 3-dimensional Z-axis fiber deposition machine
that deposits "groupings of fiber filaments" at multiple locations
normal to the plane of the composite laminate structure and cuts
each individual grouping such that a extension of each "grouping of
fiber filaments" remains above the upper skin and below the lower
skin. The preformed composite laminate then continues to be pulled
into a secondary wet-out station. Next the preformed composite
laminate travels into a pultrusion die where the extended
"groupings of fiber filaments" are all bent over above the top skin
and below the bottom skin producing a superior clinched Z-axis
fiber reinforcement as the composite laminate continues to be
pulled, catalyzed, and cured at the back section of the pultrusion
die. The composite laminate continues to be pulled by grippers that
then feed it into a gantry CNC machine that is synchronous with the
pull speed of the grippers and where computerized machining,
drilling, and cutting operations take place. This entire method is
accomplished automatically without the need for human
operators.
Inventors: |
Johnson, David W.; (San
Diego, CA) ; Hook, James M.; (Alpine, CA) ;
Garrett, Scott A.; (San Diego, CA) ; Moyers, Steven
G.; (Jamul, CA) |
Correspondence
Address: |
CHARLES C. LOGAN II
8282 UNIVERSITY AVENUE
LA MESA
CA
91941
US
|
Family ID: |
30003940 |
Appl. No.: |
10/059956 |
Filed: |
November 19, 2001 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
60281838 |
Apr 6, 2001 |
|
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60293939 |
May 29, 2001 |
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Current U.S.
Class: |
156/148 ;
156/166; 156/181 |
Current CPC
Class: |
Y10T 428/249924
20150401; Y10T 428/249923 20150401; B29C 70/088 20130101; Y10T
428/23914 20150401; E01C 9/086 20130101; Y10T 428/31504 20150401;
Y10S 411/904 20130101; E04C 2/296 20130101; Y10T 428/233 20150115;
B32B 5/02 20130101; B29C 70/24 20130101; Y10T 428/24174 20150115;
B29C 70/086 20130101; Y10T 428/249942 20150401; Y10T 29/49957
20150115 |
Class at
Publication: |
156/148 ;
156/181; 156/166 |
International
Class: |
B32B 001/00 |
Claims
What is claimed:
1. A method of clinching the top and bottom ends of Z-axis
reinforcing fibers into the respective top and bottom surfaces of a
composite laminate comprising: providing at least two rolls of
composite fiber material on at least two spools and said rolls of
composite fiber material on each of said spools having a front end;
assembling said rolls of composite fiber material into a composite
laminate preform having a top skin having a top surface, a bottom
skin having a bottom surface, an X-axis and a Y-axis; feeding said
composite laminate preform into a Z-axis fiber deposition machine
whereby Z-axis fiber bundles are deposited into said composite
laminate preform at predetermined locations along said X-axis and
said Y-axis; said Z-axis fiber bundles each having a top end that
extends a predetermined height H1 above said top surface of said
top skin and a bottom end that extends a predetermined H2 below
said bottom surface of said bottom skin; said composite laminate
preform exits said Z-axis fiber deposition machine as a modified
composite laminate preform; and feeding said modified composite
laminate preform into fastening means where said respective top and
bottom ends of said Z-axis fiber bundles are clinched into said
respective top surface of said top skin and said bottom surface of
said bottom skin of said modified composite laminate preform; said
modified composite laminate preform exits said fastening means as
an assembled composite laminate panel.
2. A method as recited in claim 1 further comprising feeding said
front ends of said rolls of composite fiber material into a resin
tank prior to assembling said rolls of composite fiber material
into said composite laminate preform; said rolls of composite fiber
material are coated with a resin in said resin tank.
3. A method as recited in claim 2 further comprising debulking said
composite laminate preform prior to exiting said resin tank.
4. A method as recited in claim 1 further comprising feeding said
composite laminate preform into a resin tank after said modified
composite laminate preform exits said Z-axis fiber deposition
machine; said modified composite laminate is coated with a resin in
said resin tank.
5. A method as recited in claim 4 wherein at least one of said
rolls of composite fiber material is in the form of a roving roll
of continuous strand mat.
6. A method as recited in claim 4 wherein at least one of said
rolls of composite fiber material is in the form of an X-Y axis
stitched fabric.
7. A method as recited in claim 4 wherein said at least one of said
rolls of composite fiber material is in the form of a woven roving
fabric.
8. A method as recited in claim 4 further comprising assembling a
core material between at least two of said rolls of composite fiber
material as they are being assembled into said composite laminate
preform.
9. A method as recited in claim 8 wherein said core material is
urethane.
10. A method as recited in claim 8 wherein said core material is a
PVC foam.
11. A method as recited in claim 8 wherein said core material is a
foam having a density in the range of 2 pounds per cubic foot to 12
pounds per cubic foot.
12. A method as recited in claim 8 wherein said core material is
balsa wood.
13. A method as recited in claim 12 wherein said balsa wood has a
density in the range of 4 lbs per cubic foot to 16 lbs per cubic
foot.
14. A method as recited in claim 4 wherein said fastening means
comprises a pultrusion die that presses said top ends of said
Z-axis fiber bundles into said top surface of said top skin and
also presses said bottom ends of said Z-axis fiber bundles into
said bottom surface of said bottom skin during the operation of
forming said assembled composite laminate panel.
15. A method as recited in claim 14 wherein said pultrusion die has
means for heating said modified composite laminate preform up to a
temperature sufficient to cause catalyzation of said modified
composite laminate preform as it becomes an assembled composite
laminate panel.
16. A method as recited in claim 4 further comprising milling means
located downstream from said fastening means; said milling means
being capable of forming bolt holes, edge routing, milling and
cut-off.
17. A method as recited in claim 16 further comprising gripper
means for transporting said assembled composite laminate panel from
said fastening means to said milling means.
18. A method as recited in claim 17 wherein said milling means is a
multi-axis CNC mailing machine.
19. A method as recited in claim 17 wherein the entire operation
starting with assembling said rolls of composite fiber material
into said composite laminate preform and continuing on to said
assembled composite laminate panel exiting said milling means is
fully automated.
20. A method of clinching the top and bottom ends of Z-axis
reinforcing fibers into the respective top and bottom surfaces of a
composite laminate comprising: providing at least two rolls of
composite fiber material on at least two spools and said rolls of
composite fiber material on each of said spools having a front end;
providing a core material between at least two of said rolls of
composite fiber material; assembling said rolls of composite fiber
material and said core material into a composite laminate preform
having a front end, a top skin having a top surface, a central
core, and a bottom skin having a bottom surface, an X-axis and a
Y-axis; feeding said front end of said composite laminate preform
into a machine whereby Z-axis fiber bundles are deposited into said
composite laminate perform at predetermined locations along said
X-axis and said Y-axis; said Z-axis fiber bundles each having a top
end that extends a predetermined height H1 above said top surface
of said top skin and a bottom end that extends a predetermined H2
below said bottom surface of said bottom skin; said composite
laminate preform exits said Z-axis fiber deposition machine as a
modified composite laminate preform having a front end; adding
resin to the preform, if required, after the deposition of said
Z-axis fiber bundles; gripping said front end of said modified
composite laminate by gripping means and feeding said modified
composite laminate preform into fastening means where said
respective top and bottom ends of said Z-axis fiber bundles are
clinched into said respective top surface of said top skin and said
bottom surface of said bottom skin of said modified composite
laminate preform; said modified composite laminate preform exits
said fastening means as an assembled composite laminate panel;
heating said modified preform up to a temperature sufficient to
cause catalyzation of said modified composite laminate preform as
it becomes an assembled composite laminate panel; transporting said
assembled composite laminate panel downstream by said gripper means
to a milling means capable of forming bolt holes, edge routing,
milling and cut-off; milling predetermined structure in said
assembled composite laminate panel with a multi-axis CNC milling
machine; and providing automation whereby the entire operating
starting with assembling said rolls of composite fiber material and
said core material into said composite laminate preform and
continuing on to said assembled composite laminate panel exiting
said milling means, is fully automated.
21. A composite laminate as produced by said method recited in
claim 17.
22. A composite laminate as recited in claim 21 having the form of
a runway panel that would be used as temporary runway matting for
military aircraft.
Description
REFERENCE TO RELATED APPLICATION
[0001] This application claims the priority of provisional patent
application 60/281838 filed on Apr. 6, 2001 and provisional patent
application, 60/293939 filed on May 29, 2001. This application also
claims the priority of U.S. patent application Ser. No. 09/922,053
filed Aug. 2, 2001.
TECHNICAL FIELD
[0002] The present invention relates to an improvement in the field
of composite laminate structures known as sandwich structures
formed with outside skins of a polymer matrix composite and an
internal core of either foam, end-grain balsa wood, or honeycomb,
and more specifically to the field of these sandwich structures
which additionally have some type of Z-axis fiber reinforcement
through the composite laminate and normal to the plane of the
polymer matrix composite skins.
BACKGROUND ART
[0003] There is extensive use in the transportation industry of
composite laminate structures due to their lightweight and
attractive performance. These industries include aerospace, marine,
rail, and land-based vehicular. The composite laminate structures
are made primarily from skins of a polymer matrix fiber composite,
where the matrix is either a thermoset or thermoplastic resin and
the fiber is formed from groupings of fiber filaments of glass,
carbon, aramid, or the like. The core is formed from end-grain
balsa wood, honeycomb of metallic foil or aramid paper, or of a
wide variety of urethane, PVC, or phenolic foams, or the like.
[0004] Typical failures in laminate structure can result from core
failure under compressive forces or in shear or, most commonly,
from a failure of the bond or adhesive capability between the core
and the composite skins (also known as face sheets). Other
failures, depending on loading may include crimpling of one or both
skins, bending failure of the laminate structure, or failure of the
edge attachment means from which certain loads are transferred to
the laminate structure.
[0005] Certain patents have been granted for an art of introducing
reinforcements that are normal to the planes of the skins, or at
angles to the normal (perpendicular) direction. This is sometimes
called the "Z" direction as it is common to refer to the
coordinates of the laminate skins as falling in a plane that
includes the X and Y coordinates. Thus the X and Y coordinates are
sometimes referred to as two-dimensional composite or 2-D
composite. This is especially appropriate as the skins are many
times made up of fiber fabrics that are stitched or woven and each
one is laid on top of each other forming plies or layers of a
composite in a 2-D fashion. Once cured these 2-D layers are 2-D
laminates and when failure occurs in this cured composite, the
layers typically fail and this is known as interlaminar
failure.
[0006] The patents that have been granted that introduce
reinforcements that are normal to the X and Y plane, or in the
generally Z-direction, are said to be introducing reinforcements in
the third dimension or are 3-D reinforcements. The purpose of the
3-D reinforcement is to improve the physical performance of the
sandwich structure by their presence, generally improving all of
the failure mechanisms outlined earlier, and some by a wide margin.
For example, we have shown that the compressive strength of a foam
core laminate structure with glass and vinyl ester cured skins can
be as low as 30 psi. By adding 16 3-D reinforcements per square
inch, that compressive strength can exceed 2500 psi. This is an 83
times improvement.
[0007] Childress in U.S. Pat. No. 5,935,680, Boyce et al in U.S.
Pat. No. 5,741,574 as well as Boyce et al in U.S. Pat. No.
5,624,622 describe Z-directional reinforcements that are deposited
in foam by an initial process and then secondarily placed between
plies of fiber fabric and through heat and pressure, the foam
crushes or partially crushes forcing the reinforcements into the
skin. Practically, these reinforcements are pins or rods and
require a certain stiffness to be forced into the skin or face
layers. Although Boyce et al describes "tow members" as the
Z-directional reinforcement, practically, these are cured tow
members, or partially cured tow members that have stiffness. As
Boyce et al describes in U.S. Pat. No. 5,624,622, compressing the
foam core will "drive" the tow members into the face sheets. This
cannot be possible unless the Z-directional or 3-D reinforcements
are cured composite or metallic pins.
[0008] A standard roll of fiberglass roving from Owens Corning,
typically comes in various yields (of yards per pound weight) and a
yield of 113 would contain on a roll or doft 40 lbs. of 113 yield
rovings. In the uncured state, these rovings are multiple filaments
of glass fiber, each with a diameter of less than 0.0005 inches.
The roving, uncured as it comes from Owens Corning, is sometimes
called a "tow", contains hundreds of these extremely small diameter
filaments. These hundreds of filaments shall be referred to as a
"grouping of fiber filaments." These groupings of fiber filaments
can sometimes be referred to, by those skilled in the art, as tows.
It is impossible to drive a virgin glass fiber tow, or grouping of
fiber filaments, as it is shipped from a glass manufacturer such as
Owens Coming, through a face sheet. The grouping of fiber filaments
will bend and kink and not be driven from the foam carrier into the
skin or face sheets as described by Boyce et al. Therefore, the
"tow" described by Boyce et al must be a rigid pin or rod in order
for the process to work as described. It will be shown that the
present invention allows easily for the deposition of these
groupings of fiber filaments, completely through the skin-core-skin
laminate structure, a new improvement in this field of 3-D
reinforced laminate structures.
[0009] This issue is further verified by an earlier patent of Boyce
et al, U.S. Pat. No. 4,808,461, in which the following statement is
made: "The material of the reinforcing elements preferably has
sufficient rigidity to penetrate the composite structure without
buckling and may be an elemental material such as aluminum, boron,
graphite, titanium, or tungsten." This particular referenced patent
depends upon the core being a "thermally decomposable material".
Other US Patents that are included herein by reference are: Boyce
et al, U.S. Pat. No. 5,186,776; Boyce et al U.S. Pat. No.
5,667,859; Campbell et al U.S. Pat. No. 5,827,383; Campbell et al
U.S. Pat. No. 5,789,061; Fusco et al. 5,589,051.
[0010] None of the referenced patents indicate that the referenced
processes can be automatic and synchronous with pultrusion, nor do
they state that the processes could be synchronous and in-line with
pultrusion. Day describes in U.S. Pat. Nos. 5,589,243 and 5,834,082
a process to make a combination foam and uncured glass fabric core
that is later molded. The glass fiber in the core never penetrates
the skins of the laminate and instead fillets are suggested at the
interface of the interior fiber fabric and the skins to create a
larger resin fillet. This is a poor way to attempt to tie the core
to the skins, as the fillet will be significantly weaker than if
the interior fiber penetrated the skins. Day has the same problem
that Boyce et al have as discussed earlier. That is, the interior
uncured fabric in Day's patent is limp and cannot be "driven" into
the skins or face sheets without being rigid. Thus the only way to
take preinstalled reinforcements in foam, and then later mold these
to face sheets under pressure, and further have the interior fiber
forced into the skins, is to have rigid reinforcements, such as
rigid pins or rods or, as in Day's case, rigid sheets.
[0011] Boyce et al in U.S. Pat. No. 5,186,776 depends on ultrasound
to insert a fiber through a solid laminate that is not a sandwich
structure. This would only be possible with a thermoplastic
composite that is already cured and certain weaknesses develop from
remelting a thermoplastic matrix after the first solidification.
Ultrasound is not a requirement of the instant invention as new and
improved means for depositing groups of fiber filaments are
disclosed. U.S. Pat No. 5,869,165 describes "barbed" 3-D
reinforcements to help prevent pullout. The instant invention has
superior performance in that the 3-D groups of fiber filaments are
extended beyond the skins on both sides of the composite laminate,
such that a riveting, or clinching, of the ends of the filaments
occurs when the ends of the filaments are entered into the
pultrusion die and cured "on-the-fly." The clinching provides
improved pull-out performance, much in the same way as a metallic
rivet in sheet metal, that is clinched or bent over on the ends,
improves the "pull-out" of that rivet versus a pin or a bonded pin
in sheet metal. This is different from the current
state-of-the-art. Fiber through the core is either terminated at
the skins, unable to penetrate the skins, or as pure rods
penetrates part or all of the skin, but is not riveted or clinched.
And many of the techniques referenced will not work with cores that
don't crush like foam. For example, the instant invention will also
work with a core such as balsa wood, which will not crush and thus
cannot "drive" cured rods or pins into a skin or face sheet.
Furthermore, the difficult, transition from a composite laminate
structure to an edge can easily be accommodated with the instant
invention. As will be shown later, a composite laminate structure
can be pultruded with clinched 3-D groupings of fiber filaments and
at the same time the edges of the pultruded composite laminate can
consist of solid composite with the same type and quantity of 3-D
grouping of fiber filaments penetrating the entire skin-central
composite-skin interface. As will be shown, the skins can remain
continuous and the interior foam can transition to solid composite
laminate without interrupting the pultrusion process.
[0012] It is an object of this invention to provide a low cost
alternative to the current approaches such that the composite
laminate structure can find its way into many transportation
applications that are cost sensitive. All prior art processes
referenced have a degree of manual labor involved and have been
only successful to date where aerospace is willing to pay the costs
for this manual labor. The instant invention is fully automatic and
thus will have extremely low selling prices. For example, earlier
it was mentioned that by adding a certain number of groups of fiber
filaments to a foam core composite laminate that the compressive
strength improved from 30 psi to over 2500 psi. This can be
achieved for only $0.30 per square foot cost. None of the existing
processing techniques referenced can compare to that
performance-to-cost ratio. This can be achieved due the automated
method of forming the composite laminate structure. Other
differences and improvements will become apparent as further
descriptions of the instant invention are given.
SUMMARY OF INVENTION
[0013] The method and apparatus for forming an improved pultruded
and clinched Z-axis fiber reinforced composite structure starts
with a plurality of upper and lower spools that supply raw material
fibers that are formed respectively into upper and lower skins that
are fed into a primary wet-out station within a resin tank. A core
material is fed into the primary wet-out station between the
respective upper and lower skins to form a composite laminate
preform. The upper and lower skins and the core are pulled
automatically through tooling where the skin material is wetted-out
with resin and the entire composite laminate is preformed in nearly
its final thickness. The composite laminate preform continues to be
pulled into an automatic 3-dimensional Z-axis fiber deposition
machine that deposits "groupings of fiber filaments" at multiple
locations normal to the plane of the composite laminate structure
and cuts individual groups such that an extension of each "grouping
of fiber filaments" remains above the upper skin and below the
lower skin.
[0014] The preformed composite laminate then continues to be pulled
into a secondary wet-out station. Next the preformed composite
laminate is pulled through a pultrusion die where the extended
"groupings of fiber filaments" are all bent over above the top skin
and below the bottom skin producing a superior clinched Z-axis
fiber reinforcement as the composite laminate continues to be
pulled, catalyzed and cured at a back section of the pultrusion
die. The composite laminate continues to be pulled by grippers that
then feed it into a gantry CNC machine that is synchronous with the
pull speed of the grippers and where computerized machining,
drilling and cutting operations take place. The entire process is
accomplished automatically without the need for human
operators.
[0015] It is an object of the invention to provide a novel improved
composite laminate structure that has riveted or clinched 3-D
groupings of fiber filaments as part of the structure to provide
improved resistance to delaminating of the skins or delaminating of
the skins to core structure.
[0016] It is also an object of the invention to provide a novel
method of forming the composite laminate structure wherein an
automatic synchronous pultrusion process is utilized, having raw
material, for example glass fabric such as woven roving or stitched
glass along with resin and core material pulled in at the front of
a pultrusion line and then an automatic deposition station places
3-D Z-axis groupings of fiber filaments through a nearly net-shape
sandwich preform and intentionally leaves these groupings longer
than the thickness of the sandwich structure, with an extra egress.
This is then followed by an additional wet-out station to
compliment an earlier wet-out station. The preform then is pulled
into a pultrusion die and is cured on the fly and the 3-D Z-axis
groupings of fiber filaments are riveted, or clinched, in the die
to provide a superior reinforcement over the prior art. The cured
composite laminate structure is then fed into a traveling CNC work
center where final fabrication! machining operations, milling,
drilling, and cut-off occur. This entire operation is achieved with
no human intervention.
[0017] It is another object of the invention to utilize core
materials that do not require dissolving or crushing as previous
prior art methods require.
[0018] It is a further an object of the invention to provide a
novel pultruded panel that can be continuous in length, capable of
100 feet in length or more and with widths as great as 12 feet or
more.
[0019] It is an additional object of the invention to produce a 3-D
Z-axis reinforced composite laminate structure wherein the edges
are solid 3-D composite to allow forming of an attachment shape or
the machining of a connection.
[0020] It is another object of the invention to provide a preferred
embodiment of a temporary runway, taxiway, or ramp for military
aircraft. This composite laminate structure would replace current
heavier aluminum structure, (known as matting) and could easily be
deployed and assembled. The 3-D Z-axis reinforcements ensure the
panels can withstand the full weight of aircraft tire loads, yet be
light enough for easy handling.
DESCRIPTION OF THE DRAWINGS
[0021] FIG. 1 is a schematic illustration of a method and apparatus
for forming continuously and automatically the subject 3-D Z-axis
reinforced composite laminate structure;
[0022] FIG. 2 is schematic vertical cross sectional view of a
pultruded composite laminate panel in a preferred embodiment, in
which the clinched 3-D Z-axis fibers have been cured on the fly,
showing side details. This panel would be used as a new lightweight
matting surface for temporary military aircraft runway use;
[0023] FIG. 3 is a magnified view taken along lines 3-3 of FIG.
2;
[0024] FIG. 4 is a magnified view taken along lines 4-4 of FIG.
3.
[0025] FIG. 5 is a schematic vertical cross-sectional view of the
pultruded sandwich panel of the preferred embodiment, just prior to
entering the pultrusion die, wherein the 3D Z-axis groupings of
fiber filaments have been deposited and they are prepared for
clinching and riveting in the die;
[0026] FIG. 6 is a magnified view taken along lines 6-6 of FIG.
5;
[0027] FIG. 7 is a magnified view taken along lines 7-7 of FIG. 6;
and
[0028] FIG. 8 is a magnified view taken along lines 8-8 of FIG.
2.
DESCRIPTION OF PREFERRED EMBODIMENT
[0029] FIG. 1 illustrates a method and application for forming a
pultruded and clinched 3-D Z-axis fiber reinforced composite
laminate structure. The pultrusion direction is from left-to-right
in FIG. 1 as shown by the arrows. The key components of the
apparatus will become evident through the following
description.
[0030] Shown in FIG. 1 are the grippers 34 and 35. These are
typically hydraulically actuated devices that can grip a completely
cured composite laminate panel 32 as it exits pultrusion die 26.
These grippers operate in a hand-over-hand method. When gripper 34
is clamped to the panel 32, it moves a programmed speed in the
direction of the pultrusion, pulling the cured panel 32 from the
die 26. Gripper 35 waits until the gripper 34 has completed its
full stroke and then takes over.
[0031] Upstream of these grippers, the raw materials are pulled
into the die in the following manner. It should be recognized that
all of the raw material is virgin material as it arrives from
various manufacturers at the far left of FIG. 1. The fiber 20 can
be glass fiber, either in roving rolls with continuous strand mat
or it can be fabric such as x-y stitched fabric or woven roving.
Besides glass, it can be carbon or aramid or other reinforcing
fiber. A core material 22 is fed into the initial forming of the
sandwich preform. The skins of the sandwich will be formed from the
layers of fiber 20 on both the top and bottom of the sandwich
preform 30. The core 22 will be the central section of the
sandwich. The core can be made of urethane or PVC foam, or other
similar foams in densities from 2 lbs. per cubic foot to higher
densities approaching 12 lbs. per cubic foot. Alternatively core 22
could be made of end-grain balsa wood having the properties of 6
lb. per cubic foot density to 16 lb. per cubic foot.
[0032] The raw materials are directed, automatically, in the
process to a guidance system in which resin from a commercial
source 21 is directed to a primary wet-out station within resin
tank 23. The wetted out preform 30 exits the resin tank and its
debulking station in a debulked condition, such that the thickness
of the panel section 30 is very nearly the final thickness of the
ultimate composite laminate. These panels can be any thickness from
0.25 inches to 4 inches, or more. The panels can be any width from
4 inches wide to 144 inches wide, or more. Preform 30 is then
directed to the Z-axis fiber deposition machine 24 that provides
the deposition of 3-D Z-axis groupings of fiber filaments. The
details as to how Z-axis filter deposition machine 24 functions is
the subject of the referenced provisional patent application
60/293,939 and U.S. patent application Ser. No. 09/922,053 filed
Aug. 2, 2001 is incorporated into this patent application by
reference. This system is computer controlled so that a wide
variety of insertions can be made. Machine 24 can operate while
stationary or can move synchronously with the gripper 34 speed.
Groupings of fiber filaments are installed automatically by this
machine into the preform 31 that is then pulled from the Z-axis
fiber deposition machine 24. Preform 31 has been changed from the
preform 30 by only the deposition of 3-D Z-axis groupings of fiber
filaments, an of which are virgin filaments as they have arrived
from the manufacturer, such as Owens Corning.
[0033] Modified preform 31 of FIG. 1 now automatically enters a
secondary wet-out station 39. Station 39 can be the primary
wet-out, eliminating station 23, as an alternative method. This
station helps in the completion of the full resin wet-out of the
composite laminate structure, including the 3-D Z-axis groupings of
fiber filaments. Preform 31 then enters pultrusion die 26 mentioned
earlier and through heat preform 31 is brought up in temperature
sufficiently to cause catalyzation of the composite laminate panel.
Exiting die 26 is the final cured panel section 32 which is now
structurally strong enough to be gripped by the grippers 34 and
35.
[0034] The sandwich structure of FIG. 1 can then be made any length
practicable by handling and shipping requirements. Downstream of
the grippers 34 and 35, the preform 32 is actually being "pushed"
into the downstream milling machine system, 36 and 37. Here a
multi-axis CNC machine (computer numerical control) moves on a
gantry synchronous with the gripper pull speed, and can machine
details into the composite laminate structure/panel on the fly.
These can be boltholes, edge routing, milling, or cut-off. The
machine 36 is the multi-axis head controlled by the computer 37.
After cut-off, the part 33 is removed for assembly or palletizing
and shipping.
[0035] FIG. 2 illustrates a vertical cross-section of one preferred
embodiment. It is a cross-section of a panel 40 that is 1.5 inches
thick and 48 inches wide and it will be used as a temporary
runway/taxiway/ or ramp for military aircraft. In remote locations,
airfields must be erected quickly and be lightweight for
transporting by air and handling. Panel 40 of FIG. 2 achieves these
goals. Because it has been reinforced with the Z-axis groupings of
fiber filaments, the panel can withstand the weight of aircraft
tires, as well as heavy machinery. Since panel 40 is lightweight,
at approximately 3 lbs. per square foot, it achieves a goal for the
military, in terms of transportation and handling. Because 40 is
pultruded automatically by the process illustrated in FIG. 1, it
can be produced at an affordable price for the military. Also shown
in FIG. 2 are edge connections, 41 and 42. These are identical but
reversed. These allow the runway panels 40 also known as matting,
to be connected and locked in place. Clearly, other applications
for these composite structures exist beyond this one
embodiment.
[0036] FIG. 3 is a magnified view taken along lines 3-3 of FIG. 2.
FIG. 3 shows the cross section of the composite laminate structure,
including the upper and lower skins 51a and 51b respectfully. Core
52, which is shown as foam, clearly could be other core material
such as end-grain balsa wood. Also shown are the several 3-D Z-axis
groupings of fiber filaments 53, which are spaced in this
embodiment every 0.25 inches apart and are approximately 0.080
inches in diameter. It can be seen from FIG. 3 that the groupings
of fiber filaments 53 are clinched, or riveted to the outside of
the skins, 51a and 51b. FIG. 4 is a magnified view taken along
lines 4-4 of FIG. 3. FIG. 4 shows core material 52 and the upper
skin section 51a and lower skin section 51b. These skin sections
are approximately 0.125 inches thick in this embodiment and
consists of 6 layers of X-Y stitched glass material at 24 oz. per
square yard weight. The Z-axis groupings of fiber filaments 53 can
be clearly seen in FIG. 4. The clinching or riveting of these
filaments, which lock the skin and core together, can clearly be
seen.
[0037] FIGS. 2, 3, and 4 show the runway matting material as it
would be produced in the method and apparatus of FIG. 1. The
schematic section 40 in FIG. 2 is fully cured as it would be
leaving pultrusion die 26. Similar drawings of these same sections
are shown for the preform of the runway matting material as it
would look just prior to entering pultrusion die 26 by FIGS. 5, 6,
and 7. FIGS. 5, 6 and 7 correlate with the preform 31 of FIG. 1.
FIGS. 2, 3, and 4 correlate with the perform 32 and the part 33 of
FIG. 1.
[0038] FIG. 5 schematically illustrates the entire matting panel 61
as a preform. The end of the panel 62 does not show the details 42,
of FIG. 2 for clarity. The lines 6-6 indicate a magnified section
that is shown in FIG. 6.
[0039] FIG. 6 shows the skins 71a and 71b, the core 72 and the 3-D
groupings of Z-axis fiber filaments 73. One can see the egressing
of the fiber filaments above and below skins 71a and 71b by a
distance H1 and H2, respectively. The lines 7-7 indicate a further
magnification which is illustrated in FIG. 7.
[0040] FIG. 7 shows the preform with the core 72 and upper skin
material 71a and a single group of Z-axis fiber filaments 73. Note
the egressed position of the fiber filaments, which after entering
the pultrusion die will be bent over and riveted, or clinched, to
the composite skin. Because the skins 71a and 71b are made of X-Y
material and the grouping of fiber filaments are in the normal
direction to X-Y, or the Z-direction, the composite skin in the
region of the 3-D grouping of fiber filaments is said to be a three
dimensional composite.
[0041] FIG. 8 is a magnified view taken along lines 8-8 of FIG. 2
and schematically depicts a core material 87, a skin material 88a
and 88b and a new interior composite material 89. As stated this
material 89 would consist of X-Y fiber material that is the same as
the skin material 88a and 88b but is narrow in width, say 2 to 3
inches wide in this matting embodiment. The 3-D groupings of Z-axis
fiber filaments 84 are deposited by the newly developed Z-axis
deposition machine 24 in FIG. 1, and are operated independent of
the density of the material. The 3-D groupings of fiber Z-axis
filaments can be easily deposited through either the core material
87 or the higher density X-Y material 89. The interlocking
connecting joint 85 can be either machined into the shape of 85 in
FIG. 8 or can be pultruded and shaped by the pultrusion die. In
FIG. 8 joint 85 is machined. If it were pultruded, the 3-D
groupings of Z-axis fiber filaments in 85 would show riveted or
clinched ends. Clearly other interlocking joints or overlaps could
be used to connect matting panels.
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